High Reynolds Number Tests of the NASA SC(2)-0012 Airfoil in the Langley 0.3-Meter Transonic ...

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NASA Technical Memorandum 89 102
High Reynolds Number Tests of
the NASA SC(2)-0012 Airfoil
in the Langley 0.3-Meter
Transonic Cryogenic Tunnel
Raymond E. Mineck and Pierce L. Law-ing
JULY 1987
NASA Technical Memorandum 89 102
High Reynolds Number Tests of
the NASA SC(2)-0012 Airfoil
in the Langley 0.3-Meter
Transonic Cryogenic Tunnel
Raymond E. Mineck and Pierce L. Lawing
LangZey Research Center
Hampton, Virginia
National Aeronautics
and Space Administration
Scientific and Technical
Information Office
1987
SUMMARY
A wind-tunnel investigation of the NASA SC(2)-0012 airfoil has been conducted
in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This supercritical airfoil sec-
tion is 12 percent thick and symmetrical. This investigation supplements the two-
1 dimensional airfoil studies of the Advanced Technology Airfoil Test program. Tests
were conducted at various combinations of stagnation temperature and pressure to
cover a Mach number range from 0.60 to 0.84 and a Reynolds number range from 6 to
40 x 106 based on a 6.0-in. (15.24-cm) airfoil chord. The angle of attack was
varied from -4.0' to 4.5'.
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No corrections for wind-tunnel wall interference have been made to the data.
The aerodynamic results are presented as pressure distributions and integrated force
and moment coefficients without any analysis.
I INTRODUCTION
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The National Aeronautics and Space Administration (NASA) has conducted a system-
atic study of well-known conventional airfoils and advanced technology airfoils over
a wide range of Reynolds number. This study, described in detail in reference 1 and
referred to as the Advanced Technology Airfoil Tests (ATAT) program, was carried out
in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The program was
divided into two phases: a correlation phase involving conventional and advanced
airfoils tested in other facilities and an advanced technology phase involving new
' airfoils without a significant data base. One of the objectives of the program was
to expand the airfoil data base at high Reynolds numbers.
One of the airfoils tested in the correlation phase of the ATAT program was part
of a family of supercritical airfoils developed by NASA. Details of the design phi-
losophy used to develop the family of airfoils are presented in reference 2. This
airfoil was designed for application to a transport-type airplane wing, and the
results are presented in reference 3. The family also included a symmetric super-
critical airfoil for nonlifting applications such as the vertical tail on a
transport-type airplane. A model of this airfoil was built to demonstrate a novel
model fabrication technique described in reference 4. This model was tested in the
0.3-m TCT to determine i f there were any limitations associated with the fabrication
technique when the model was subjected to load at cryogenic temperatures. The
results would add high Reynolds number data for a symmetric supercritical airfoil to
the ATAT airfoil data base of cambered supercritical airfoils.
The model tests of the symmetric supercritical airfoil were conducted in the
8-in. by 24-in. (20-cm by 60-cm) two-dimensional test section of the 0.3-m TCT.
Reference 5 describes the operating characteristics and operating envelope of the
tunnel. Tests were conducted at various combinations of stagnation temperature and
pressure to cover a Mach number range from 0.60 to 0.84 and a Reynolds number range
from 6 to 40 X lo6 based on the airfoil chord. Integrated forces and moments and a
summary of the airfoil chordwise pressure distributions are presented at selected
values of Mach number and Reynolds number without analysis.
SYMBOLS
The measurements and calculations were made in U.S. Customary Units. These
measurements have been converted to the International System of Units (SI) which are
shown in parentheses. Factors relating the two systems of units are presented in
reference 6.
b model span, 8.0 in. (20.32 cm)
section drag coefficient measured on tunnel centerline, positive downstream
‘d
cm section pitching moment, resolved about quarter-chord point, positive
nose up
section normal-force coefficient, positive in z-direction
Cn
local pressure coefficient
cP
C model chord, 6.0 in. (15.24 an)
M, f ree-s tream Mach number
free-stream Reynolds number based on model chord
RC
X chordwise position from leading edge of model (positive aft) , in. (cm)
Y spanwise position on model from centerline (positive to right looking
upstream), in. (cm)
z vertical distance from model chord plane (positive measured up), in. (cm)
a section geometric angle of attack (positive leading edge up), deg
standard deviation in Mach number during a data recording
OM
standard deviation in Reynolds number during a data recording
OR
WIND TUNNEL AND MODEL
Wind Tunnel
The tests were conducted in the 8-in. by 24-in. (20-cm by 60-cm) two-dimensional,
slotted-wall test section of the 0.3-m TCT. A schematic diagram of the tunnel show-
ing the various components is presented in figure 1. A photograph of the tunnel is
presented in figure 2. The 0.3-m TCT is a continuous-operation, fan-driven, cryo-
genic pressure tunnel that uses gaseous nitrogen as a test medium. The tunnel is
capable of operating at stagnation temperatures from just above the boiling point of
liquid nitrogen (LN2), approximately 80 K, to just above room temperature, approxi-
mately 327 K, and a t stagnation pressures from 1.2 to 6.0 atm. The test-section Mach
number has been calibrated from 0.20 to about 0.92. This combination of test condi-
tions provides Reynolds numbers up to 55 X lo6 based on a model chord of 6.0 in.
(15.24 cm). Additional details about the tunnel may be found in reference 5.
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A photograph of a model in the test section is presented in figure 3. The top
of the plenum and the test-section top wall (ceiling) have been removed to show the
model installation. The two-dimensional test section has solid sidewalls and a
slotted top w a l l (ceiling) and bottom wall (floor). There are two slots in the top
and bottom slotted walls with pressure orifices located on the centerline on the slat
between the slots. A photograph of the model mounted between the turntables just
prior to installation in the test section is presented in figure 4, and a sketch of
the test section is presented in figure 5.
A computer-driven mechanism controls the angular position of the turntables to
set the angle of attack (AOA). The mechanism has a traversing range from -2OO to
20°, which can be offset from 0' in either direction during the model installation.
It is driven by an electric stepper motor, which is connected through a yoke to the
perimeter of both turntables. This arrangement drives both ends of the model through
the angle-of-attack range to eliminate possible model twisting. The angular position
of the turntables and, therefore, the geometric angle of attack of the model are
recorded using the output of a digital shaft encoder that is geared to one of the
turntables.
A computer-controlled survey mechanism traverses a momentum rake vertically to
measure the total pressure loss in the model wake. The mechanism provides two mount-
ing locations on the left sidewall for the rake: either at tunnel station 8.3 in.
(21.0 cm) or at 10.2 in. (26.0 cm). For this test, the wake survey measurements were
made at the rear station, which placed the measurement plane about 1.2 chord lengths
downstream of the airfoil trailing edge. The survey mechanism is driven by an elec-
tric stepper motor and is designed to translate the rake at speeds from about
0.1 in/sec (0.25 cm/sec) to about 6 in/sec (15 cm/sec). It has a total traversing
range of 10 in. (25.4 cm). The stroke (that portion of the total traversing range
used in a given survey), the number of points in a survey, the number of samples at
each point in the survey, and the speed of the survey mechanism can be controlled by
the operator in the control room to meet the research requirements. The vertical
position of the rake is recorded using the output from a digital shaft encoder geared
to the survey mechanism.
The momentum rake has nine total pressure probes as shown in figure 6. The six
active probes used in this test were located at the following spanwise stations:
y/(b/2) = 0 (centerline), -0.125, -0.250, -0.375, -0.500, and -0.750.
Model
The model used in this test was a 12-percent-thick, symmetric, advanced tech-
nology supercritical airfoil designed by NASA for a Reynolds number of 30 X lo6.
The model chord is 6.0 in. (15.24 cm). The design and measured model ordinates are
presented in table I. It should be noted that the airfoil design technique was
applied up to the 60-percent-chord station. The rest of the airfoil was defined as
a straight line to the trailing edge. The model was fabricated using a novel process
described below; details are available in reference 4.
The model was designed as two mating halves with channels routed into the mating
surfaces to provide a path between tubing leading to the pressure measurement system
and the pressure orifice. A photograph of the model at this stage in the fabrication
is shown in figure 7. A t one end of the channel, the pressure orifices are pre-
drilled outward from the mating surface at the proper angle to intersect the model
surface after the final machining. A t the opposite end, tubes are brazed in to form
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a connection with the wind-tunnel pressure instrumentation system. The photograph
also shows the channels matching on both surfaces to form pressure orifices at the
trailing edge. The fan-shaped channel configurations are the top and bottom leading-
edge pressure orifice rows.
After the machining on the mating surfaces is complete, the two halves of the
model are assembled with brazing foil between the mating surfaces and with the con-
nector tubes in place. The model is then placed in a vacuum-brazing oven. Weights
are placed on the model to provide a moderate pressure to force the two pieces
together as the brazing foil melts. The model can then be rough machined and the
integrity of the pressure orifices verified before the final machining is done.
This fabrication process provided much freedom in selecting the location of the
pressure orifices. The orifices were arranged as shown in figure 8. This report
presents the results from the chordwise rows staggered near the centerline on the
upper and lower surfaces. There were 25 orifices on the upper surface and 27 on the
lower surface. All pressure orifices were 0.010 in. (0.25 mm) in diameter except for
the short chordwise row of orifices to the right of the centerline which were
0.005 in. (0.13 mm) in diameter. A listing of the orifice locations is presented
in table 11.
TEST INSTRUMENTATION AND PROCEDURES
Test Instrumentation
A detailed discussion of the instrumentation and procedures for the calibration
and control of the 0.3-m TCT can be found in reference 7. For two-dimensional air-
foil tests, the 0.3-m TCT is equipped to obtain static pressure measurements on the
airfoil model surface, total pressure measurements in the model wake, and static
pressure measurements on the sidewalls, top wall, and bottom wall of the test section.
Except for the w a l l pressures, all measurements use individual pressure transducers.
Tunnel test conditions.- The tunnel test conditions were determined by three
primary measurements: the total pressure, the static pressure, and the total tem-
perature. The total pressure and static pressure were measured by individual quartz
differential pressure transducers referenced to a vacuum. The transducer has a range
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from -100 to t100 psid (-6.8 to +6.8 atm) It has an accuracy of 20.006 psid
(t0.0004 atm) plus 20.012 percent of the pressure reading. The stagnation tempera-
ture was measured by a platinum resistance thermometer. The analog output from these
devices was converted by individual digital voltmeters. Further details of this
instrumentation are presented in reference 7.
Airfoil model pressures.- The pressures on the airfoil model are measured by
individual transducers connected by tubing to each orifice on the model. The trans-
ducers are of a commercially available, high-precision, variable-capacitance type.
The maximum range of these differential transducers is from -100 to +lo0 psid
(-6.8 to +6.8 atm) with an accuracy of t0.25 percent of the reading from -25 to
+lo0 percent of full scale. They are located external to the tunnel and its high-
pressure cryogenic environment, but as close as possible to the test section to
minimize the tubing length and reduce the response time. To provide increased accu-
racy, the transducers are mounted both on thermostatically controlled heater bases to
maintain a constant temperature and on "Shock" mounts to reduce possible vibration
effects. The electrical signals from the transducers are connected to individual
signal conditioners located in the tunnel control room. The signal conditioners are
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autoranging and have seven ranges available. As a result of the autoranging capa-
bility, the analog output to the data acquisition system is kept at a high level even
though the pressure transducer may be operating at the low end of its range.
Wall pressures.- The tunnel floor and ceiling pressures are measured using a
scanivalve system capable of operating ten 48-port scanivalves. Because of the
large changes in the pressure of the tunnel over its operational range, the same
capacitance-type pressure transducers and autoranging signal conditioners described
above are used in the scanivalve instead of the more typical strain gauge transducer.
These pressures are normally used in the wall correction procedure and are not pre-
sented in this report.
Wake pressures.- The total pressure loss in the model wake is measured with the
momentum rake described previously. The pressure in each of the six operational
tubes is measured with the same type of pressure transducer described above. The
static pressures on the right sidewall at nine vertical positions at the tunnel sta-
tion opposite the momentum rake are measured with nine individual pressure trans-
ducers. The transducers used are of the same capacitance type described above but
with a maximum range from -20 to +20 psid (-1.36 to +1.36 atm).
Procedures
Figure 9 shows the test program (+ versus %) used in this investigation.
For these tests, no transition strip was placed on the model because of the high
Reynolds number capability of the 0.3-m TCT. The selection of test conditions was
made in an effort to overlap experimental and theoretical work for some of the air-
foils tested in the ATAT.program.
The following procedure was used to set the test conditions. The tunnel total
pressure and temperature and the fan speed were set for the desired Mach number and
Reynolds number. The angle of attack was set for near-zero model l i f t and then the
data were recorded. The angle of attack was then increased to the next angle, the
tunnel conditions were adjusted, and the data were recorded. If the angle exceeded
the desired value when setting the angle of attack, the angle was reduced well below
the desired angle before attempting to reset the angle again. This is a standard
testing technique to minimize the possibility of hysteresis effects. This process
was repeated up to the maximum positive angle of 4.5'. The angle of attack was then
returned to near-zero model l i f t and the process was repeated for increasing negative
angles of attack.
At the beginning of the data recording process, the computer surveys the model
wake and presents the results graphically to the operator. The operator then chooses
the stroke (the upper and lower traversing limits) for this data point so that about
80 percent of the points are within the wake. The rake is then positioned to the
beginning of the stroke near the ceiling. Twenty samples of the airfoil static
pressures, the test conditions, the momentum rake total pressures, the wake static
pressures, and the wall pressures on the scanivalve were recorded over a period of
1 sec. Since there were individual transducers for each orifice on the model, each
sample consisted of simultaneous static pressure readings from all orifices on the
model. These 20 samples were used to compute the test conditions, the normal force,
and the pitching moment. The wake survey mechanism was synchronized with the scani-
valves so that the rake was moved to the next vertical location when the scanivalves
advanced to a new port. After the appropriate settling time, another 20 samples were
recorded at the new rake location. This procedure continued until the scanivalves
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completed their stepping at which time the rake continued to step at a predetermined
rate through the remaining portion of the wake. For this test, the number of steps
within the stroke was held constant at 50.
DATA ACQUISITION, REDUCTION, AND QUALITY
Data Acquisition
Data were recorded on magnetic tape with a computer-controlled, high-speed,
digital data acquisition system located in the control room of the 0.3-m TCT. This
system has a total of 192 analog channels with 8 selectable ranges from +4 to f131 mv
and a resolution of 1 part in 8191. In addition, there are 16 digital input channels.
A l l analog data were filtered with a 10-Hz low-pass filter. An operating and acqui-
sition program is used by the computer to scan the data acquisition hardware and to
write the raw data on tape.
Data Reduction
The test Mach number is based on an average of the Mach number distributions
measured on the turntables during the calibration of the "empty" test section.
Because the tunnel operating envelope includes high pressures and low temperatures,
real-gas effects are included in the data reduction for the tunnel test conditions
using the thermodynamic properties of nitrogen gas calculated from the Beattie-
Bridgeman equation of state. This equation of state has been shown in reference 8 to
give essentially the same thermodynamic properties and flow calculation results in
the temperature-pressure regime of the 0.3-m TCT as those given by the more compli-
cated Jacobsen equation of state. Detailed discussions of real-gas effects when
testing in cryogenic nitrogen are contained in references 9 and 10.
Section normal-force and pitching-moment coefficients are calculated using the
first 20 samples from numerical integration (based on the trapezoidal method) of the
local surface pressure coefficient measured at each orifice multiplied by an appro-
priate weighting factor (incremental area).
The section drag coefficient is calculated from the wake survey pressures by
first computing an incremental or point drag coefficient by the method of refer-
ence 11 for each rake tube pressure at each rake position. These point drag coeffi-
cients are then numerically integrated across the model wake, again based on the
trapezoidal method. Specifically, the point drag coefficients are compared one by
one to a "threshold" value of drag coefficient which accounts for a nonzero pressure
decrement outside the model wake. Based on experience from previous tests, the
threshold value was set a t 0.0002 for this test. If, in the integrating process, the
individual coefficient is greater than or equal to the threshold value, the weighting
factor (incremental area) is applied and the incremental drag is included in the
running sum of the total drag. If the individual coefficient is less than the
threshold, the weighting factor is set equal to zero and the incremental drag is
not included in the running sum of the total drag. The results of this integration
are total drag coefficients at each of the six momentum rake tube locations. The
data reduction program then provides a correction which subtracts that summed portion
of the individual incremental drag coefficients within the wake that is attributable
to the threshold level. These corrected values are the ones used for the discussion
of all the drag data. In this report, all drag data pertain to the tube on the
tunnel centerline.
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Data Quality
In all wind-tunnel testing, and especially in transonic testing, the steadiness
of the tunnel flow conditions, such as Mach number, have direct bearing on the qual-
ity of the final aerodynamic data. With the use of individual pressure transducers
on each of the model pressure orifices, and with all the model data being recorded
at the time of the first rake step, Mach number fluctuations in the surface pressure,
normal force, and pitching-moment data for any given point are essentially nonexis-
tent. The possibility of Mach number fluctuations during the time required for the
rake to complete the survey of the wake was checked for several angles of attack for
a single set of tunnel test conditions, M,,, = 0.60 and 0.80 at Rc = 6, 15, and
40 X lo6 at zero lift. The standard deviations in Mach number UM and in Reynolds
number oR during the wake surveys are given, respectively, as follows:
,
Standard deviation in Mach number
OM
-
for Reynolds numbers of
M,,,
6 X lo6 15 x lo6 40 X lo6
0.60 0.0009 0.0006 0.0011
.80 .0029 .0014 .0014
6 X lo6 15 X lo6 40 X lo6
I
0.011 0.012 0.070
.015 .021 .099
The tunnel sidewalls induce an error in Mach number; the slotted top and bottom walls
can induce an error in angle of attack and Mach number because of the model l i f t and
blockage. No corrections for wall effects have been applied to the experimental data
presented in this report.
PRESENTATION OF RESULTS
The results are presented in the following order:
Figure
..........................................
Effect of M , on integrated forces 10
..........................................
Effect of R, on integrated forces 11
Effect of a on chordwise pressure distribution:.
...............................................................
Rc = 6 X lo6 ............................................................... 12
R, = 9 x 106 13
..............................................................
Rc = 15 X lo6 .............................................................. 14
R, = 30 X lo6 .............................................................. 15
R, = 40 X lo6 16
............................
Effect of M, on chordwise pressure distribution 17
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DISCUSSION
The effect of Mach number on the integrated force and moment coefficients is
presented in figure 10. Zero l i f t did not coincide with 0' angle of attack. The
difference may be attributed to a misalignment between the surface used to set Oo
angle of attack and the model chord plane, to the difference between the design and
measured model ordinates, or to flow angularity.
The effect of chord Reynolds number on the integrated force and moment coeffi-
cients is presented in figure 11. There was only a small effect of chord Reynolds
number on the model normal-force and pitching-moment results. There was a more sig-
nificant effect on the drag, especially at the lower chord Reynolds number.
The effect of angle of attack on the chordwise pressure distributions is pre-
sented in figures 12 to 16. The plots have been arranged so that corresponding
positive and negative normal-force coefficients are side by side. The pressure dis-
tributions are consistent; that is, the upper surface pressure distribution at a
positive normal force is virtually the same as the lower surface pressure distribu-
tion at the same negative normal-force coefficient. The effect of Mach number on the
chordwise pressure distribution is presented in figure 17 for near-zero normal-force
coefficient.
The model was inspected after testing was completed. No mechanical problems
such as debonding of the model halves and warping of the model surface were detected.
This inspection, along with the consistent pressure distributions described above,
demonstrate that the novel model fabrication technique is acceptable for cryogenic
wind- tunnel models.
CONCLUDING REMARKS
A wind-tunnel investigation of the NASA SC(2)-0012 airfoil has been conducted in
the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation supplements the
NASA/U.S. industry two-dimensional airfoil studies of the Advanced Technology Airfoil
Test program. This investigation was designed to demonstrate a novel model fabrica-
tion technique and to test a NASA advanced technology airfoil from low to flight-
equivalent Reynolds numbers.
Both objectives of this investigation were met. The results are presented in
graphical form without corrections.
NASA Langley Research Center
Hapt on, VA 2 366 5- 5 2 2 5
May 1, 1987
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